Cold nozzle optimisation

Abstract

The present disclosure relates to optimisation of a cold nozzle, or bypass exit area, for a gas turbine engine, in particular for a geared turbofan gas turbine engine. Example embodiments include a method of optimising a geared turbofan gas turbine engine for an aircraft, the method comprising: determining expected service parameters for the aircraft, the expected service parameters including an expected range of travel for the aircraft; selecting components for the geared turbofan gas turbine engine to define a first smaller cold nozzle area if the range of travel is within a first smaller range and to define a second larger cold nozzle area if the range of travel is within a second larger range.

We claim: 
     
         1 . A method of optimising a geared turbofan gas turbine engine for an aircraft, the method comprising:
 determining expected service parameters for the aircraft, the expected service parameters including an expected range of travel for the aircraft;   selecting components for the geared turbofan gas turbine engine to define a first smaller cold nozzle area if the range of travel is within a first smaller range and to define a second larger cold nozzle area if the range of travel is within a second larger range.   
     
     
         2 . The method of  claim 1 , wherein the geared turbofan gas turbine engine comprises an engine core and a nacelle defining a bypass duct and the cold nozzle area. 
     
     
         3 . The method of  claim 2 , wherein the components of the geared turbofan gas turbine engine include:
 the nacelle, or a portion thereof; and/or an outer casing of the engine core, or a portion thereof.   
     
     
         4 . The method of  claim 3 , wherein the step of selecting components for the geared turbofan gas turbine engine from two alternative selections defining respective first and second cold nozzle areas. 
     
     
         5 . The method of  claim 1 , wherein a specific thrust of the geared turbofan gas turbine engine with the first smaller cold nozzle area is at least 1% greater than a specific thrust of the geared turbofan gas turbine with the second larger cold nozzle area. 
     
     
         6 . The method of  claim 5 , wherein the specific thrust of the geared turbofan gas turbine engine with the first smaller cold nozzle is between 1% and 3% greater than the specific thrust of the geared turbofan gas turbine with the second larger cold nozzle. 
     
     
         7 . The method of  claim 5 , wherein the specific thrust of the geared turbofan gas turbine engine is within the range of 80 to 110 Nkg −1 s. 
     
     
         8 . The method of  claim 1 , wherein a fan inlet area of the geared turbofan gas turbine engine is between 4.1 and 11.2 m 2 . 
     
     
         9 . The method of  claim 1 , wherein the geared turbofan gas turbine engine is configured to have an fan tip air angle of between 57 and 62 degrees at cruise, the fan tip air angle θ being defined as: 
       
         
           
             
               θ 
               = 
               
                 
                   tan 
                   
                     - 
                     1 
                   
                 
                  
                 
                   ( 
                   
                     
                       V 
                       ThetaBladeTip 
                     
                     
                       Vx 
                       air 
                     
                   
                   ) 
                 
               
             
           
         
         
           
             
               
                 
                   where 
                    
                   
                       
                   
                    
                   
                     V 
                     ThtaBladeTip 
                   
                 
                 = 
                 
                    
                   
                     ω 
                     · 
                     
                       D 
                       2 
                     
                   
                    
                 
               
               ; 
             
           
         
       
       ω is the fan rotational speed in radians/second, D is the outer tip diameter of the fan in metres at its leading edge and Vx air  is the mean axial velocity of the flow into the fan over the leading edge of the fan blades. 
     
     
         10 . The method of  claim 1 , further comprising assembling the geared turbofan gas turbine engine. 
     
     
         11 . The method of  claim 1 , wherein the first smaller range of travel is between 556 and 1296 km (300 and 700 nautical miles) and the second larger range of travel is greater than 1296 km (700 nautical miles). 
     
     
         12 . A plurality of gas turbine engines for a plurality of aircraft, each gas turbine engine comprising:
 an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;   a fan located upstream of the engine core, the fan comprising a plurality of fan blades;   a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and   an outer housing defining a bypass duct between the outer housing and the engine core and further defining a cold nozzle,   wherein each of a first subset of the plurality of gas turbine engines comprises components defining a first smaller cold nozzle area and each of a second subset of the plurality of gas turbine engines comprises components defining a second larger cold nozzle.   
     
     
         13 . The plurality of gas turbine engines of  claim 12 , wherein the components defining the first and second cold nozzle areas are the outer housing, or portion thereof, and/or an outer casing, or portion thereof, of the engine core. 
     
     
         14 . The plurality of gas turbine engines of  claim 12 , wherein the first plurality of gas turbine engines are for aircraft having a first smaller range of travel and the second plurality of gas turbine engines are for aircraft having a second larger range of travel. 
     
     
         15 . The plurality of gas turbine engines of  claim 14 , wherein the first smaller range of travel is between 556 and 1296 km (300 and 700 nautical miles) and the second larger range of travel is greater than 1296 km (700 nautical miles). 
     
     
         16 . The plurality of gas turbine engines of  claim 12 , wherein each gas turbine engine ( 10 ) has a nominally identical engine core, fan and gearbox. 
     
     
         17 . The plurality of gas turbine engines according to  claim 12 , wherein for each gas turbine engine:
 the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;   the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and   the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.